Aircraft autonomous pushback

ABSTRACT

The invention provides methods and systems for controlling speed of an aircraft during an autonomous pushback manoeuvre, i.e. under the aircraft&#39;s own power without a pushback tractor. The method includes applying a torque to at least one landing gear wheel of the aircraft, the torque being in a direction opposite to the backwards rolling direction of rotation of the landing gear wheel. The torque applied does not exceed a limit for ensuring aircraft longitudinal stability. For longitudinal stability the torque applied should not cause the aircraft to risk a tip-over event.

RELATED APPLICATIONS

The present application is a continuation of U.S. application Ser. No.14/912,692 filed Feb. 18, 2016, which is a National Phase ofInternational Application Number PCT/GB2014/052217 filed Jul. 21, 2014,and claims priority from Great Britain Application No. 1315012.3, filedAug. 22, 2013. The disclosures of all of the above-listed priorapplications are hereby incorporated by reference herein in theirentirety.

FIELD OF THE INVENTION

The present invention relates to reversing of an aircraft on the groundusing an autonomous taxiing system, and in particular the inventionrelates to decelerating the aircraft when reversing.

BACKGROUND OF THE INVENTION

It is sometimes necessary to move an aircraft in reverse whilst theaircraft is on the ground. For example aircraft are frequently pushedbackwards away from an airport gate, a so called “pushback” manoeuvre,by a vehicle known as a pushback tractor or tug. The same tractor or tugcan also move the aircraft forward if desired. Whilst some aircraft havethe capability to reverse under their own power using the main aircraftengines, e.g. by using reverse thrust in a manoeuvre known as “powerback”, this is not permitted for civil aircraft as the jet or prop washfrom the engines can cause damage to nearby terminal buildings orfacilities.

Recently it has been proposed to equip aircraft with an autonomous wheeldrive taxi system that can drive one or more of the aircraft landinggear wheels in rotation so that the aircraft can taxi under its ownpower. The ability to drive the aircraft landing gear wheel in reverseusing the wheel drive system enables an autonomous pushback operationwithout the main engines running, as well as a forward taxiing operationwith or without one or more of the main engines running. The autonomouspushback operation makes the autonomous wheel drive taxi systemparticularly suitable for civil aircraft but the system has wideapplicability to a wide variety of aircraft including: civil andmilitary; fixed wing, rotary wing and powered lift; manned and unmanned,etc.

For conventional pushback operations, braking is performed by thetractor. Use of the brake pedals is generally prohibited as there is arisk of damaging the nose landing gear and the tractor. With anautonomous wheel drive taxi system, i.e. without a tractor unit, brakingwill need to be performed autonomously on the aircraft. Conventionalaircraft braking systems are typically not designed to perform thisfunction, being more suited to high energy dissipation in the landingphase and low energy dissipation working against the thrust of the mainaircraft engines in the taxi phase. Conventional aircraft brakingsystems are therefore generally unsuited to braking during an autonomouspushback operation. The autonomous pushback introduces the risk ofaircraft tip over (where the aircraft tends to rotate nose up about itspitch axis, possibly causing a tail strike) and aircraft runaway (whereif the autonomous pushback manoeuvre is performed on a slope thelongitudinal component of the gravity acceleration can become higherthan the rolling resistance of the tyres on the ground such that theaircraft accelerates above the desired pushback speed).

SUMMARY OF THE INVENTION

A first aspect of the invention provides a method of controlling speedof an aircraft during backwards motion of the aircraft when in contactwith the ground, the method comprising: applying a torque to at leastone landing gear wheel of the aircraft, the torque being in a directionopposite to the backwards rolling direction of rotation of the landinggear wheel, wherein the torque applied does not exceed a limit forensuring aircraft longitudinal stability.

A further aspect of the invention provides an autonomous pushbackbraking system for an aircraft having a wheel drive system for drivingone or more of the aircraft's landing gear wheels in rotation, whereinthe wheel drive system is operable to drive the wheel in rotation toeffect backwards motion of the aircraft when in contact with the ground,and a means for applying a torque to at least one landing gear wheel ofthe aircraft, the torque being in a direction opposite to the backwardsrolling direction of rotation of the landing gear wheel, and wherein thetorque applied does not exceed a limit for ensuring aircraftlongitudinal stability.

For longitudinal stability the torque applied to decelerate the aircraftduring a pushback manoeuvre should not cause the aircraft to risk atip-over event. Accordingly the torque limit may be selected such thata) the aircraft cannot tip back onto its tail, and/or b) a nose landinggear of the aircraft does not part contact with the ground, and/or c) asubstantially vertical load through a nose landing gear wheel does notfall below a threshold at which a steering centring device maintains thewheel of the nose landing gear straight.

The torque limit may be based upon one or more of the followinginstantaneous aircraft parameters when the torque is applied: a slopeangle of the ground over which the aircraft is moving; the centre ofgravity of the aircraft; the mass of the aircraft; an aircraft inertiamoment around the aircraft lateral (y) axis; the backwards speed of theaircraft.

The backwards motion of the aircraft may be effected by a wheel actuatorcarried by the aircraft for driving one or more of the aircraft'slanding gear wheels in rotation and/or by gravity due to a slope angleof the ground over which the aircraft is moving.

In an autonomous pushback manoeuvre the backwards motion of the aircraftis effected autonomously in the absence of an external tractor unit.However, pushback runaway may also be caused by an inadvertent releaseof the park brake even when a pushback manoeuvre has not been commanded.

The step of applying torque to the landing gear wheel may compriseapplying a braking torque to the wheel using a friction brake assembly.

The aircraft speed may be measured and if this speed exceeds apredetermined limit at which the aircraft longitudinal stability cannotbe ensured then an indication is displayed in the aircraft cockpit.

The braking torque may be effected by a braking control system or by apark brake system. The braking torque may be initiated by a pilot input,e.g. using cockpit brake pedals or a park brake lever.

The braking control system may send a braking command to only a limitednumber of braking wheels. For example where the aircraft has N brakingwheels the braking control system may send a braking command to a numbern of the braking wheels where n<N.

The braking control system may limit the maximum braking clampingpressure applicable to the friction brake assemblies to no more than alimit at which the aircraft longitudinal stability is ensured.

The maximum braking clamping pressure may be variable depending on themass and longitudinal centre of gravity position of the aircraft.

The braking control system may implement a braking law that commandsinitially a low brake pressure which rises with increasing time.

The braking torque may be applied whilst the wheel actuator is drivingthe aircraft backwards.

The wheel actuator torque and the braking torque may be controlled by acommon controller. The controller may receive input of the aircraftspeed and control the wheel actuator torque and the braking torquetowards a target speed.

The step of applying torque to the landing gear wheel may alternativelycomprise applying a braking torque to the wheel using a generator.

The generator may be coupled either to an electrical network of theaircraft or to a resistor for dissipating the electrical energygenerated by the generator.

The generator is preferably a motor/generator used to drive one or moreof the aircraft's landing gear wheels in rotation to effect thebackwards motion of the aircraft.

The motor/generator may be selectively coupled to the landing gearwheel(s) by a drive path. Preferably the drive path includes a gearmounted to the wheel rim and a pinion, wherein the pinion is moveablebetween an engaged position in which the pinion is in driving engagementwith the wheel gear and a disengaged position in which the pinion isphysically separated from the wheel gear.

BRIEF DESCRIPTION OF THE DRAWINGS

Embodiments of the invention will now be described with reference to theaccompanying drawings, in which:

FIG. 1 illustrates an aircraft performing an autonomous pushbackoperation;

FIG. 2 illustrates the aircraft on sloping ground;

FIG. 3 illustrates a wheel drive system of the aircraft;

FIG. 4 illustrates a plot of the main contributors to the longitudinalstability and performance of the aircraft during an autonomous pushbackoperation;

FIG. 5 illustrates a scheme for braking the aircraft using theaircraft's braking control system;

FIG. 6 illustrates a scheme for braking the aircraft using theaircraft's park brake;

FIG. 7 illustrates a ramp braking law of the aircraft braking controlsystem;

FIG. 8 illustrates a control scheme for controlling the wheel drivesystem and the braking control system of the aircraft;

FIG. 9 illustrates a scheme for dissipating electrical power generatedby a generator to the aircraft electrical power network; and

FIG. 10 illustrates a scheme for dissipating electrical power generatedby a generator to a resistor.

DETAILED DESCRIPTION OF EMBODIMENT(S)

FIG. 1 shows an aircraft 1 having a fuselage 2 including a nose 2 a anda tail 2 b, wings 3, main engines 4, nose landing gear 5 and mainlanding gear 6. The aircraft has two main landing gears 6, one on eitherside of the aircraft centreline, and a single nose landing gear 5forming a tripod. Each landing gear 5, 6 has a diablo configuration withtwo wheels.

The aircraft 1 is typical of a short range single aisle passenger jetaircraft, although it will be appreciated that the invention hasapplicability to a wide variety of aircraft types as mentioned above. Inparticular the aircraft may have a greater or fewer number of landinggears; and each landing gear may have any number of wheels, includingone.

Each main landing gear 6 has a wheel drive system 10, shown in detail inFIG. 3. The wheel drive system 10 is for driving one wheel of the mainlanding gear 6 (typically the outboard wheel but may alternatively bethe inboard wheel) in rotation to taxi the aircraft on the ground. Thewheel drive system 10 is operated without the main engines 4 running forreversing the aircraft. The wheel drive system 10 is operated eitherwith or without one or more of the main engines running for taxiing theaircraft forward. For the avoidance of doubt the forward direction isalong the aircraft longitudinal axis nose first and the reversedirection is along the aircraft longitudinal axis tail first.

In FIG. 1 the aircraft is shown reversing, indicated by the directionarrow R, whilst the landing gear wheels are in contact with the ground Gwhich is substantially level, i.e. zero slope with the horizontal. InFIG. 2 the aircraft 1 is shown reversing, indicated by the directionarrow R, whilst the landing gear wheels are in contact with the ground Gwhich has a slope angle alpha (a) to the horizontal, h.

Wheel Drive System

FIG. 3 shows a partial view of the wheel drive system 10. The mainlanding gear 6 includes a telescopic shock-absorbing main leg 12,including an upper telescopic part (main fitting) and a lower telescopicpart (slider) 13. The upper telescopic part is attached to the aircraftfuselage or wing (not shown) by its upper end (not shown). The lowertelescopic part supports an axle 14 carrying a pair of wheels 16, one oneither side of the main leg. Each wheel comprises a tyre supported by ahub 18 (only the hub 18 of one wheel 16 is shown in FIG. 3, forclarity). The wheels 16 are arranged to rotate about the axle 14 toenable ground movement of the aircraft, such as taxiing or landing.

Each wheel hub 18 has a rim 19 for holding the tyre (not shown). Thewheel drive system 10 includes a driven gear 20 attached to the hub 18so as to be rotatable with the wheel 16, the driven gear 20 comprising aroller gear 34 formed by two rigid annular rings 35 connected togetherby a series of rollers 36 extending around the rings to form acontinuous track. The rollers 36 are each rotatable about a pin (notshown) which extends between the annular rings 35 to form a rigidconnection between the annular rings 35. One of the annular rings 35comprises a plurality of connection extension tabs 37 which provide arigid connection to the hub 18.

The wheel drive system 10 further comprises a wheel actuator 50comprising a motor 52 which rotates an output sprocket 60 (drive pinion)via a gearbox 70. The sprocket 60 is a wheel-type sprocket withradially-extending teeth which can interlock with the rollers 32 of theroller gear 34.

The wheel actuator 50 is supported by a bracket which is rigidlyconnected to the axle 14 of the landing gear and pivotally connected tothe motor 52 about a pivot axis. The wheel actuator 50 may alternativelybe mounted on the upper telescopic part (main fitting) or lowertelescopic part 13 (slider). A linear actuator 58, such as adirect-drive roller-screw electro-mechanical linear actuator, extendsbetween the bracket 56 (at an end nearest the axle 14) and the motor 52.Thus, linear movement of the actuator 58 is translated to rotationalmovement of the wheel actuator 50 causing the sprocket 60 to movebetween an engaged position in which the sprocket teeth interlock withthe rollers 32 of the roller gear 34, and a disengaged position in whichthe sprocket teeth are physically separated from the rollers 32 of theroller gear 34. The sprocket 60 is therefore moved in a substantiallyradial direction with respect to the roller gear 34 axis of rotationbetween the engaged and disengaged positions.

It will be appreciated that the wheel drive system 10 may take a varietyof forms. The wheel drive system 10 illustrated is an example of anopen-geared arrangement where the engagement/disengagement of the wheeldrive system is by moving the drive pinion in a substantially radialdirection into/out of positive driving engagement with the driven gear.The drive pinion and driven gear respectively may be formed as asprocket and roller gear (as illustrated); a sprocket and roller chain;a roller gear and sprocket; a roller chain and sprocket; or toothedgears, e.g. spur gears. Alternatively the drive pinion may move in asubstantially axial direction (along the axis of rotation of the drivepinion) into and out of driving engagement with the driven gear.

Yet further alternatively a clutch device may be provided between themotor and the driven wheel. The motor may be disposed within the wheelhub or mounted adjacent the wheel. The driven wheel is permanentlyengaged with a portion of a drive path between the motor and the drivenwheel, and the clutch device may make or break the drive path betweenthe motor and the driven wheel. The engagement between the drive pathand the driven wheel may be a geared drive or may be a friction drive.

Although the figures only show features of the wheel drive system 10 fordriving one of the wheels 16, these features be mirrored for the otherwheel 16. That is, one wheel drive system 10 may be provided for eachwheel 16. For a landing gear 10 with four or more wheels 16, a wheeldrive system 10 may be provided for each of the wheels 16, or for onlytwo of them. In other embodiments it may be possible to have one motor52 shared between two wheel drive systems 10. That is, the motor 52 maybe arranged to rotate the output sprockets of each drive system.Additionally or alternatively a wheel drive system may drive one or morewheels of the nose landing gear 5.

Inside the wheel hub 18 is a friction brake arrangement indicatedgenerally at 40. The friction brake arrangement 40 may be ofconventional type and so will not be described in detail here.Generally, however, the aircraft friction brake arrangement includes astator part and a rotor part comprising a stack of carbon disks. A brakeactuator, which may be hydraulically or electrically operated, buildsthe pressure on the carbon stack to convert rotational torque of thewheel into heat thereby decelerating the aircraft. The brake actuator(s)are controlled by a braking control system (BCS) of the aircraft. TheBCS responds to inputs, e.g. pilot and autopilot inputs, and commands abrake pressure accordingly.

Aircraft Longitudinal Stability

Whilst the tripod arrangement of landing gears 5, 6 is generally stablein the aircraft longitudinal direction when the aircraft is travellingforwards on the ground and the aircraft is decelerated by the frictionbrakes of the main landing gear, studies have shown that when theaircraft is performing an autonomous pushback operation application onlya low level of braking torque using the friction brakes may besufficient to disturb the aircraft longitudinal stability and risk a tipover event.

Pushback Tip Over

For longitudinal stability three different tip over cases have beenidentified:

-   -   Tail strike—The aircraft nose lifts and the tail impacts on the        ground.    -   Nose lift up—The wheels of the nose landing gear momentarily        lose contact with the ground but the aircraft does not sit on        the tail.    -   Steering cam engagement—The load on the nose landing gear        steering centring cam goes below a predetermined level at which        the centring cam engages to maintain the nose landing gear        wheels inline (zero degree position facing forwards along the        aircraft centreline).

FIG. 4 illustrates the effect of particular aircraft parameters on theaircraft longitudinal stability (x-axis) and aircraft performance(y-axis). The parameters considered are:

-   -   Aircraft Mass    -   Aircraft longitudinal centre of gravity (CG X) position    -   Apron (ground) slope    -   Inertia moment about pitch (y) axis, Iyy    -   Autonomous Pushback Speed    -   Brake Gain    -   Braking Rise time (time to build brake pressure to commanded        level)

It has been identified that of these parameters a combination of an aftCG position, with a high pushback speed and a high apron slope angle maypose the greatest risk of a tip over event during an autonomous pushbackmanoeuvre. Although not shown in FIG. 4, it should also be noted thatwith an aft CG a higher mass is also more detrimental to the aircraftlongitudinal stability.

Pushback Runaway

Simulations have also been performed for the pushback runaway conditionif the autonomous pushback is performed on a slope. The longitudinalcomponent of the gravity acceleration adds a component of the gravityforce in the direction of the motion. The runaway condition isencountered when the gravity component on the aircraft longitudinal axisis higher than the sum of the aerodynamic drag and the friction forceson the ground. The runaway phenomenon is not affected by the aircraftmass. The highest slope angle achievable at the gate is assumed to be1.15 degrees, or 2%. The study has shown if the autonomous pushback wasto be performed without restrictions on the taxiway slope, the aircraftwould need to be provided with a means of applying a relatively lowretardation force to protect from pushback runaway.

Design Solution to Address Pushback Tip Over and/or Pushback Runaway

Various design solutions have been identified and these will bedescribed detail.

1. Pushback Speed Limitation with Pilot Braking

One solution is to limit the pushback speed. By selecting the worst casecombination of parameters (Brake Gain, Slope, Mass, CG, . . . ) a speedlimit can be identified for a particular aircraft type up to which theaircraft BCS can apply full brake pressure without risks of tip over.Whilst this solution would negate risks of tip over the speed limit islikely to be very low, e.g. around 1 knot. This may be considered toolow a pushback speed to be commercially viable.

As shown in FIG. 5 the aircraft speed is detected, e.g. by a wheel speedsensor 100 (or alternatively from the aircraft inertial referencesensing system, a GPS system or a resolver in the wheel drive system10), and if this is judged to be greater than or equal to the speedlimit 102 then a cockpit indication 104 alerts the pilot that thereversing speed limit has been reached. Upon receiving the indication,e.g. as a visual, audible or tactile warning, the pilot would berequired to react by using the conventional foot brake pedals 106. TheBCS 108 would interpret the braking input in the usual manner andcommand a brake pressure to the wheel brake assemblies 110 depending onthe degree of brake pedal deflection. Since the autonomous taxi speedlimit is selected such that any brake pressure up to the full brakepressure will not risk a tip over event the pilot can make any brakeinput to stop the aircraft.

The components of conventional aircraft BCS generally have a significantlevel of inaccuracy, particularly at low brake pressures. For example,in a conventional hydraulic aircraft braking system the zero torquepressure (ZTP), i.e. the braking system pressure that will apply zerobraking torque, may have a nominal value of around 15-20 bar. However,the actual ZTP may be significantly lower than this nominal value, e.g.around 10 bar. Furthermore, a valve regulating the hydraulic pressure atthe brake actuators may have a tolerance of +/−5 bar or more at lowbrake pressures. There are yet further sources of inaccuracy.

Due to these BCS inaccuracies when the pilot makes a brake input tomaintain the speed below the autonomous taxi speed limit it may bedifficult for the pilot to bring the aircraft to a smooth stop. Thiswould result in an unpleasant and discontinuous pushback manoeuvre.

A partial solution would be to limit the autonomous taxi speed to ahigher speed limit than that mentioned above, and reducing the allowableMass-CG combination for the aircraft, excluding the areas of the CGenvelope where the tip over risk is higher, e.g. high mass and aft CG.These extremities of the mass-CG envelope of the aircraft would need tobe restricted accordingly. However a limitation in the CG envelope is alimitation in the flexibility of the operability of the aircraft.

2. Use of Park Brake

Another solution is to use a modified park brake to stop the aircraft.In particular it would be possible to limit the rise rate of the brakingapplication through one or more restrictors limiting the flow rate inthe hydraulic system. Simulations have been carried out to evaluate foreach speed the acceptable rise rate that would stop the aircraft, withinacceptable passenger comfort level (assumed longitudinal decelerationlower than 0.2 G). The study indicates that the brake rise time wouldincrease rapidly with increasing pushback speed.

In order to implement this solution the maximum pushback speed wouldstill need to be limited to a low speed if the brake rise time is to bekept within acceptable limits. Since the park brake rise time isgenerally applicable for all aircraft operations it is not consideredfeasible to increase the brake rise time to the extent that itsperformance is limited, e.g. for taxiing with engines or hybrid(combined wheel taxi and engine taxi) operations. Accordingly, a lowpushback speed limit would need to be imposed which may be similar tosolution 1 above. Indeed solution 1 may be preferred.

FIG. 6 illustrates an implementation of the park brake solution, whichis reliant on a cockpit indication 104 of when the pushback speed isgreater than or equal to the pushback speed limit. In response to thecockpit indication the pilot would need to activate the park brake lever112 in the cockpit which controls the park brake 116 having a brake riselimiter 114 to ensure than the deceleration is within the acceptablepassenger comfort level (assumed longitudinal deceleration lower than0.2 G) at the speed limit.

3. Braking with Limited Number of Brakes

A further solution involves commanding braking through a limited numberof brakes. In order to achieve the same aircraft level retardation forcea reduced number of brakes need to be commanded with a higher brakingpressure.

For alternate braking the braking command is sent to a wheel pair (e.g.one main landing gear wheel on each side of the aircraft), hence theminimum number of braking wheels would be two. In this case the clampingbraking pressure commanded could be raised to approximately double thatof the baseline case where all four wheels of the main landing gears 6are braked.

For normal braking a single braking command can be sent to one of themain landing gear wheels. In this case the clamping braking pressurecommanded could be raised further to approximately quadruple that of thebaseline case where all four wheels of the main landing gears 6 arebraked.

This solution alone may not be able to accommodate full braking pressurewithout any risk of tip over, but may be used in combination with one ormore other solutions presented here.

4. Single or Variable Maximum Brake Pressure Limitation

In this solution an overall brake pressure limitation is set to ensurethe braking clamping pressure is never so high as to cause aircraft tipover. Preferably such pressure limitation is customised for eachoperational point in the CG-Mass diagram to optimise the maximumpressure applicable by the system depending on the aircraft loadcondition. By accurately measuring the vertical load through eachlanding gear leg (the sprung mass), e.g. using strain gauges on the mainstrut, and with prior knowledge of the un-sprung mass of each landinggear, the CG X position and mass of the aircraft can be determined. Byplotting these values on the CG-Mass diagram the maximum brake clampingpressure can be determined.

5. Ramp Braking Law

Analysing the dynamic response of the aircraft in the limit case fornose lift-up we observe that there is a nose lift-up ‘peak’ immediatelyafter the brake application, where the nose landing gear shock absorbersextend and the pitch angle reaches a maximum. Subsequently the aircrafttends to put the nose back down due to its weight.

This solution aims to prevent nose lift-up and obtain an optimisedperformance during braking by implementing a braking law that commandsinitially the maximum pressure that ensures no nose lift-up, and afterthe nose lift-up ‘peak’ increases gradually the braking command, up tothe maximum system pressure.

The increase in pressure is initially very gentle and subsequently veryrapid once the nose lift-up peak is safely overcome. Such behaviour canbe modelled with an exponential function, e.g.:

$P = {\frac{P_{2} - {P_{1}e^{\tau}}}{1 - e^{\tau}} - {\frac{P_{1} - P_{2}}{1 - e^{\tau}}e^{\frac{t\; \tau}{T}}}}$

FIG. 7 illustrates graphically how the system braking pressure varieswith time, where P2 is the final pressure (max system pressure), P1 isthe initial pressure (max initial pressure to ensure no nose lift-up),and T is the total time from brake application to full brakes applied.The parameter, τ, regulates the behaviour of the ramp law. A high valueof τ makes flatter the first part of the ramp law and steeper the secondpart, and vice versa.

This solution may be preferred, as it guarantees stability withoutcompromising on performance (in all cases full braking applicationwithin T seconds), but has to take into account the limitationsintroduced by the low accuracy of conventional braking systems.

6. Braking Against the Wheel Actuator Motors

The main limitations of using the conventional braking system to providethe aircraft with retardation force include both the low limit ofbraking pressure that can be applied without risking aircraft tip-up andthe low accuracy of conventional braking systems at very low pressures.

A solution that could mitigate both limitations includes using theautonomous wheel drive taxi system in conjunction with the BCS. Theamount of wheel drive taxi system torque in the direction of motion isadditional torque that needs to be demanded to the braking system toprovide some retardation, and the inaccuracies of the BSCS can bebalanced by a fine tuning of the wheel drive taxi system torque, toensure the global retardation torque is sufficiently accurate.

This solution is particularly beneficial for the pushback runawayproblem, for which the amount of required braking torque is very low andalso very dependent on certain parameters (Slope, Mass, Brake Gain, . .. ).

FIG. 8 shows an example of a closed loop speed controller 200 forregulating the wheel drive taxi system torque during pushback. When abinary pushback command is sent from the pilot the autonomous wheeldrive taxi system (or “eTaxi”) motors 204 target the pushback speed 202,and in absence of apron slope the speed 206 is achieved and maintainedby the aircraft, and there is no need to apply braking.

In the presence of a downhill slope (in the direction of motion) theaircraft tends to accelerate, as described above, and when it overshootsthe target speed plus a defined tolerance, a threshold speed 208, a setamount of braking is commanded 210 to the BCS 212.

The additional braking torque 214 will act as an ‘additional’ resistiveforce that will tend to slow down the aircraft, hence the eTaxi speedcontroller 200 will reduce the amount of torque required to maintain thespeed, so the overall contribution of eTaxi and braking will be aretardation force that is accurately controlled by the eTaxi controller.

When the slope decreases and drops below the pushback runaway limit theeTaxi system will saturate the torque controller to match the amount ofbraking provided. When this happens the eTaxi controller 200 willcommand braking release, as it will have detected the end of the slope.

7. Braking Through the eTaxi System

The solution uses the dissipation of the power generated in the eTaxielectric motors (used as generators) 300 to provide the torque necessaryto stop the aircraft. This solution requires the capability to sinkpower back into the aircraft electrical network 302 (as shown in FIG.9), or to use a resistor 304 to dissipate the energy through the JouleEffect (as shown in FIG. 10).

Considering the maximum aircraft mass, M, and the pushback speed, V, themaximum kinetic energy, K, to be dissipated is (in the ideal case of100% motor/generator efficiency):

$K = {\frac{1}{2}{{MV}^{2}:}}$

The amount of heat dissipated through the resistor and the maximumreachable brake temperature, Tmax, sizes the minimum required mass ofthe resistor at aircraft level. Assuming three consecutive full brakesapplication, and an initial brake temperature T1, the mass of theresistor m_(res):

$m_{res} = \frac{3\; K}{C_{s}\Delta \; T}$

where Cs is the specific heat coefficient of the material chosen for theresistor and ΔT=Tmax−T1. The mechanical power to be dissipated for eachbrakes application through the resistor, assuming no loss in theelectro/mechanical chain is:

$P = \frac{K}{\Delta \; t}$

The mechanical power is the kinetic energy over the time interval Δt forthe braking, and is equal to the electric power dissipated through Jouleeffect, expressed as a function of the back emf ε induced in the motor:

$\frac{K}{\Delta \; t} = \frac{ɛ^{2}}{R}$

Assuming a resistor formed as a simple wire dissipating heat (hence notconsidering the insulation, and the resistor case) the resistance R canbe expressed through the second Ohm's Law as a function of mass of theresistor and geometry:

$R = {\frac{\rho \; l}{A} = {\frac{\rho \; {lA}}{A^{2}} = {\frac{\rho \; V}{A^{2}} = \frac{m_{res}\rho}{\delta \; A^{2}}}}}$

Where ρ is the material's resistivity, l the length of the wire in theresistor, A the wire cross section's area, V the wire's total volume,and δ the material's density. Thus we obtain:

$ɛ = {{\frac{K}{A}\sqrt{\frac{C_{m}}{\Delta \; T\; \Delta \; t}}\mspace{14mu} {with}\mspace{14mu} C_{m}} = \frac{\rho}{\delta \; C_{s}}}$

where C_(m) is constant, depending on the chosen material.

ε is the amount of back emf that the system has to be able to tolerate.In order to minimise this quantity the time interval (duration of thebraking) has to be increased, and a material providing a low Cm has tobe selected. This means to select a material with low resistivity (goodconductor), a high density and a high specific heat coefficient.Suitable candidate materials include steel, copper and aluminium, withcopper or steel being preferred.

The weight and space impact of the resistor will also need to take intoaccount a consistent amount of insulating material around the resistorwire, adding mass and volume. It may be preferable to split the amountof power to dissipate among more resistors, e.g. at least one perlanding gear, in order to optimise the dissipation within the availablespace.

For dissipation of the power generated in the electric motors the wheeldrive system 10 needs to be capable of being back-driven. The wheeldrive system 10 shown in FIG. 3 is particularly suitable as the sprocket60 is in positive meshing engagement with the roller gear 34 when thewheel drive system is in operation. Other wheel drive systems thatinclude an over-running clutch in the drive path between the wheel andthe motor would not be suitable as these only permit drive torque to bepassed from the motor to the wheel and not vice versa.

The clutch-less wheel drive system 10 shown in FIG. 3 is alsoparticularly beneficial as the physical separation between the drivepinion and the driven gear when the system is moved to the disengagedposition ensures that on landing there is no drive path between thewheel and the motor.

Although the invention has been described above with reference to one ormore preferred embodiments, it will be appreciated that various changesor modifications may be made without departing from the scope of theinvention as defined in the appended claims.

1. A method of controlling speed of an aircraft during backwards motionof the aircraft when in contact with the ground, the method comprising:applying a torque to at least one landing gear wheel of the aircraft,the torque being in a direction opposite to the backwards rollingdirection of rotation of the landing gear wheel, wherein the torqueapplied does not exceed a limit for ensuring aircraft longitudinalstability.
 2. A method according to claim 1, wherein the torque applieddoes not exceed a limit at which aircraft longitudinal stability isensured such that the aircraft cannot tip back onto its tail.
 3. Amethod according to claim 1 wherein the torque applied does not exceed alimit at which aircraft longitudinal stability is ensured such that anose landing gear of the aircraft does not part contact with the ground.4. A method according to claim 1 wherein the torque applied does notexceed a limit at which aircraft longitudinal stability is ensured suchthat a substantially vertical load through a nose landing gear wheeldoes not fall below a threshold at which a steering centring devicemaintains the wheel of the nose landing gear straight.
 5. A methodaccording to claim 1, wherein the torque limit is based upon one or moreof the following instantaneous aircraft parameters when the torque isapplied: a slope angle of the ground over which the aircraft is moving;the centre of gravity of the aircraft; the mass of the aircraft; anaircraft inertia moment around the aircraft lateral (y) axis; thebackwards speed of the aircraft.
 6. A method according to claim 1,wherein the backwards motion of the aircraft is effected by a wheelactuator carried by the aircraft for driving one or more of theaircraft's landing gear wheels in rotation.
 7. A method according toclaim 1, wherein the backwards motion of the aircraft is effectedautonomously in the absence of an external tractor unit.
 8. A methodaccording to claim 1, wherein the backwards motion of the aircraft iseffected by gravity due to a slope angle of the ground over which theaircraft is moving.
 9. A method according to claim 1, wherein the stepof applying torque to the landing gear wheel comprises applying abraking torque to the wheel using a friction brake assembly.
 10. Amethod according to claim 9 wherein the aircraft speed is measured andif this speed exceeds a predetermined limit at which the aircraftlongitudinal stability cannot be ensured then an indication is displayedin the aircraft cockpit.
 11. A method according to claim 9, wherein thebraking torque is initiated by a pilot input.
 12. A method according toclaim 9, where the braking torque is effected by a braking controlsystem.
 13. A method according to claim 9, wherein the braking torque iseffected by a park brake system.
 14. A method according to claim 12,wherein the aircraft has N braking wheels and the braking control systemsends a braking command to a number n of the braking wheels where n<N.15. A method according to claim 12, wherein the braking control systemlimits the maximum braking clamping pressure applicable to the frictionbrake assemblies to no more than a limit at which the aircraftlongitudinal stability is ensured.
 16. A method according to claim 15,wherein the maximum braking clamping pressure is variable depending onthe mass and longitudinal centre of gravity position of the aircraft.17. A method according to claim 12, wherein the braking control systemimplements a braking law that commands initially a low brake pressurewhich rises with increasing time.
 18. A method according to claim 6,wherein the braking torque is applied whilst the wheel actuator isdriving the aircraft backwards.
 19. A method according to claim 18,wherein the wheel actuator torque and the braking torque are controlledby a common controller.
 20. A method according to claim 19, wherein thecontroller receives input of the aircraft speed and controls the wheelactuator torque and the braking torque towards a target speed.
 21. Amethod according to claim 1, wherein the step of applying torque to thelanding gear wheel comprises applying a braking torque to the wheelusing a generator.
 22. A method according to claim 21, wherein thegenerator is coupled either to an electrical network of the aircraft orto a resistor for dissipating the electrical energy generated by thegenerator.
 23. A method according to claim 21, wherein the generator isa motor/generator used to drive one or more of the aircraft's landinggear wheels in rotation to effect the backwards motion of the aircraft.24. A method according to claim 23, wherein the motor/generator isselectively coupled to the landing gear wheel(s) by a drive path.
 25. Amethod according to claim 24, wherein the drive path includes a gearmounted to the wheel rim and a pinion, wherein the pinion is moveablebetween an engaged position in which the pinion is in driving engagementwith the wheel gear and a disengaged position in which the pinion isphysically separated from the wheel gear.
 26. An autonomous pushbackbraking system for an aircraft having a wheel drive system for drivingone or more of the aircraft's landing gear wheels in rotation, whereinthe wheel drive system is operable to drive the wheel in rotation toeffect backwards motion of the aircraft when in contact with the ground,and a means for applying a torque to at least one landing gear wheel ofthe aircraft, the torque being in a direction opposite to the backwardsrolling direction of rotation of the landing gear wheel, and wherein thetorque applied does not exceed a limit for ensuring aircraftlongitudinal stability.
 27. A system according to claim 26, furthercomprising a friction brake assembly for applying a braking torque tothe wheel.
 28. A system according to claim 26, further comprising asensor for determining the aircraft speed relative to the ground.
 29. Asystem according to claim 28, further comprising a cockpit indicator forindicating to a pilot when the aircraft speed exceeds a predeterminedlimit at which the aircraft longitudinal stability cannot be ensured.30. A system according to claim 27, further comprising a braking controlsystem for sending a braking command to the friction brake assembly. 31.A system according to claim 30, wherein the braking control system isadapted to receive a pilot braking input.
 32. A system according toclaim 27, further comprising a park brake system for sending a brakingcommand to the friction brake assembly.
 33. A system according to claim30, wherein the aircraft has N braking wheels and the braking controlsystem is adapted to send a braking command to a number n of the brakingwheels where n<N.
 34. A system according to claim 30, wherein thebraking control system is adapted to limit the maximum braking clampingpressure applicable to the friction brake assemblies to no more than alimit at which the aircraft longitudinal stability is ensured.
 35. Asystem according to claim 34, wherein the maximum braking clampingpressure is variable depending on the mass and longitudinal centre ofgravity position of the aircraft.
 36. A system according to claim 30,wherein the braking control system is adapted to implement a braking lawthat commands initially a low brake pressure which rises with increasingtime.
 37. A system according to claim 27, wherein the friction brakesystem is adapted to apply the braking torque whilst the wheel drivesystem is driving the aircraft backwards.
 38. A system according toclaim 37, further comprising a common controller for controlling thewheel drive system torque and the braking torque.
 39. A system accordingto claim 38, wherein the controller is adapted to receive input of theaircraft speed and to control the wheel actuator torque and the brakingtorque towards a target speed.
 40. A system method according to claim26, further comprising a generator for applying a braking torque to thewheel.
 41. A system according to claim 40, wherein the generator iscoupled either to an electrical network of the aircraft or to a resistorfor dissipating the electrical energy generated by the generator.
 42. Asystem according to claim 40, where in the generator is amotor/generator forming part of the wheel drive system.
 43. A systemaccording to claim 42, wherein the motor/generator is selectivelycoupled to the landing gear wheel(s) by a drive path.
 44. A systemaccording to claim 43, wherein the drive path includes a gear mounted tothe wheel rim and a pinion, wherein the pinion is moveable between anengaged position in which the pinion is in driving engagement with thewheel gear and a disengaged position in which the pinion is physicallyseparated from the wheel gear.
 45. An aircraft including the autonomouspushback braking system of claim
 26. 46. A system according to claim 26,wherein the wheel drive system is supported by a bracket which isrigidly connected to the axle, main fitting or slider part of thelanding gear.